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GOE 681 AIRFOIL (goe681-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 681 AIRFOIL (goe681-il)
Reynolds number: 100,000
Max Cl/Cd: 44.27 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe681-il-100000.txt
Download as CSV file: xf-goe681-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 681 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.2899   0.10408   0.09954  -0.0563   0.9685   0.1707
  -8.250  -0.4519   0.06441   0.05930  -0.0764   0.9483   0.0974
  -8.000  -0.4505   0.05875   0.05340  -0.0789   0.9398   0.0955
  -7.750  -0.4615   0.05402   0.04836  -0.0783   0.9288   0.0945
  -7.500  -0.4596   0.04824   0.04201  -0.0797   0.9218   0.0939
  -7.250  -0.4636   0.04527   0.03865  -0.0770   0.9111   0.0940
  -7.000  -0.4446   0.04172   0.03455  -0.0777   0.9051   0.0945
  -6.750  -0.4372   0.03969   0.03211  -0.0755   0.8961   0.0950
  -6.500  -0.4162   0.03755   0.02951  -0.0751   0.8894   0.0958
  -6.250  -0.3805   0.03544   0.02700  -0.0769   0.8854   0.0974
  -6.000  -0.3716   0.03470   0.02625  -0.0741   0.8753   0.0988
  -5.750  -0.3389   0.03358   0.02505  -0.0751   0.8698   0.1019
  -5.500  -0.2977   0.03229   0.02340  -0.0773   0.8663   0.1069
  -5.250  -0.2922   0.03192   0.02317  -0.0739   0.8556   0.1094
  -5.000  -0.2573   0.03102   0.02218  -0.0749   0.8503   0.1160
  -4.750  -0.2130   0.02988   0.02112  -0.0775   0.8472   0.1255
  -4.500  -0.2118   0.02985   0.02107  -0.0732   0.8359   0.1318
  -4.250  -0.1765   0.02935   0.02072  -0.0742   0.8305   0.1518
  -4.000  -0.1306   0.02907   0.02044  -0.0768   0.8274   0.1855
  -3.750  -0.1356   0.02964   0.02110  -0.0715   0.8147   0.1949
  -3.500  -0.0956   0.02936   0.02069  -0.0731   0.8100   0.2226
  -3.250  -0.0408   0.02865   0.02013  -0.0771   0.8074   0.2498
  -3.000  -0.0433   0.02910   0.02047  -0.0719   0.7931   0.2626
  -2.750   0.0149   0.02807   0.01961  -0.0760   0.7893   0.2927
  -2.500   0.0326   0.02795   0.01947  -0.0740   0.7787   0.3083
  -2.250   0.0708   0.02724   0.01872  -0.0750   0.7720   0.3268
  -2.000   0.1202   0.02629   0.01770  -0.0778   0.7683   0.3455
  -1.750   0.1294   0.02644   0.01796  -0.0746   0.7568   0.3552
  -1.500   0.1713   0.02564   0.01716  -0.0761   0.7512   0.3727
  -1.250   0.2201   0.02466   0.01620  -0.0787   0.7472   0.3943
  -1.000   0.2272   0.02498   0.01651  -0.0751   0.7346   0.4081
  -0.750   0.2712   0.02404   0.01568  -0.0769   0.7294   0.4316
  -0.500   0.2914   0.02388   0.01561  -0.0752   0.7193   0.4516
  -0.250   0.3273   0.02307   0.01499  -0.0757   0.7117   0.4846
   0.000   0.3756   0.02169   0.01402  -0.0780   0.7070   0.5665
   0.250   0.4990   0.02057   0.01399  -0.0927   0.6941   0.9454
   0.500   0.6262   0.01989   0.01294  -0.1101   0.6820   0.9978
   0.750   0.6545   0.01960   0.01245  -0.1100   0.6710   1.0000
   1.000   0.6672   0.01964   0.01239  -0.1072   0.6583   1.0000
   1.250   0.6950   0.01932   0.01185  -0.1066   0.6481   1.0000
   1.500   0.7081   0.01935   0.01180  -0.1037   0.6341   1.0000
   1.750   0.7244   0.01936   0.01170  -0.1013   0.6204   1.0000
   2.000   0.7475   0.01924   0.01139  -0.1000   0.6075   1.0000
   2.250   0.7695   0.01915   0.01113  -0.0984   0.5934   1.0000
   2.500   0.7841   0.01925   0.01113  -0.0957   0.5773   1.0000
   2.750   0.8007   0.01935   0.01111  -0.0933   0.5612   1.0000
   3.000   0.8185   0.01946   0.01108  -0.0912   0.5453   1.0000
   3.250   0.8372   0.01961   0.01107  -0.0892   0.5298   1.0000
   3.500   0.8578   0.01979   0.01105  -0.0875   0.5151   1.0000
   3.750   0.8764   0.02005   0.01115  -0.0856   0.5007   1.0000
   4.000   0.8913   0.02042   0.01143  -0.0832   0.4865   1.0000
   4.250   0.9096   0.02078   0.01167  -0.0813   0.4741   1.0000
   4.500   0.9336   0.02109   0.01176  -0.0804   0.4630   1.0000
   4.750   0.9467   0.02157   0.01222  -0.0778   0.4514   1.0000
   5.000   0.9695   0.02200   0.01248  -0.0768   0.4420   1.0000
   5.250   0.9860   0.02249   0.01296  -0.0748   0.4325   1.0000
   5.500   1.0111   0.02299   0.01329  -0.0743   0.4249   1.0000
   5.750   1.0257   0.02358   0.01391  -0.0720   0.4168   1.0000
   6.000   1.0529   0.02406   0.01422  -0.0719   0.4099   1.0000
   6.250   1.0675   0.02474   0.01495  -0.0698   0.4031   1.0000
   6.500   1.0869   0.02531   0.01549  -0.0684   0.3964   1.0000
   6.750   1.1193   0.02591   0.01589  -0.0692   0.3908   1.0000
   7.000   1.1283   0.02670   0.01683  -0.0662   0.3854   1.0000
   7.250   1.1455   0.02742   0.01759  -0.0646   0.3803   1.0000
   7.500   1.1724   0.02807   0.01815  -0.0646   0.3756   1.0000
   7.750   1.1978   0.02891   0.01895  -0.0645   0.3714   1.0000
   8.000   1.2048   0.02984   0.02006  -0.0613   0.3670   1.0000
   8.250   1.2194   0.03068   0.02096  -0.0595   0.3625   1.0000
   8.500   1.2437   0.03141   0.02166  -0.0591   0.3585   1.0000
   8.750   1.2790   0.03234   0.02246  -0.0607   0.3550   1.0000
   9.000   1.2792   0.03354   0.02390  -0.0567   0.3520   1.0000
   9.250   1.2830   0.03480   0.02533  -0.0535   0.3488   1.0000
   9.500   1.2909   0.03598   0.02662  -0.0509   0.3454   1.0000
   9.750   1.3075   0.03695   0.02765  -0.0496   0.3421   1.0000
  10.000   1.3359   0.03785   0.02850  -0.0500   0.3390   1.0000
  10.250   1.3528   0.03929   0.02999  -0.0490   0.3362   1.0000
  10.500   1.3319   0.04126   0.03224  -0.0428   0.3340   1.0000
  10.750   1.3022   0.04356   0.03476  -0.0357   0.3319   1.0000
  11.000   1.2640   0.04673   0.03815  -0.0290   0.3296   1.0000
  12.000   1.3045   0.05376   0.04539  -0.0242   0.3192   1.0000
  12.250   0.6415   0.14087   0.13330  -0.0417   0.3581   1.0000
  12.500   0.6737   0.14398   0.13643  -0.0417   0.3560   1.0000
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